A common configuration for a helicopter includes a first rotor, often called the main rotor or the lift rotor, rotating in a generally horizontal plane above the helicopter airframe and a second rotor, often called the tail rotor or anti-torque rotor, mounted on a tail boom and rotating in a generally vertical plane oriented to produce a sideways thrust in the direction of yaw. The pitch of the tail rotor blades, i.e., the angle between the chord line of the blade profile and the direction of rotation of the tail rotor, can be varied so as to increase or decrease the amount of sideways thrust produced by the tail rotor. The sideways thrust of the tail rotor serves three related purposes: first, since the tail rotor is located on a tail boom a distance from the main rotor, its sideways thrust produces a moment which serves to offset the torque produced on the airframe of the helicopter by the rotation of the main rotor blade; second, the sideways thrust of the tail rotor provides yaw axis control for the helicopter; and third, the sideways thrust of the tail rotor may work in conjunction with sideways thrust of the main rotor when the helicopter is translating laterally through the air.
The total sideways thrust produced by the tail rotor is known as the tail rotor authority. Factors affecting the total authority produced by a tail rotor include blade size and profile, rotational speed, angle of attack of the tail rotor blades, the pitch of the tail rotor blades, and the air density. The angle of attack is the angle between the chord line of the blade profile and the "relative wind", i.e., the direction at which the air approaches the tail rotor blade. This angle of attack is affected by the rotor blade pitch, the direction of travel of the helicopter and the presence of cross winds. A cross wind which reduces the angle of attack reduces the overall authority produced by the tail rotor, diminishing the control available to the pilot. The pitch is the angle between the chord line of the blade profile and the direction of blade rotation. The pitch is not affected by cross winds. The pilot controls the pitch of the tail rotor blades through the use of control pedals. Increasing the blade pitch results in greater tail rotor authority and decreasing the blade pitch results in less tail rotor authority. Air density also affects the tail rotor authority. Other factors being equal, the greater the air density, the greater the authority produced by the tail rotor, and similarly, the lower the air density, the less authority produced by the tail rotor. Since air density is difficult to measure directly, it is often calculated using equations based on the "standard atmosphere." The U.S. Standard Atmosphere is described in MARK's STANDARD HANDBOOK FOR MECHANICAL ENGINEERS, p. 11-67 (E. A. Avallone et al. eds., 9th ed. 1978), and Table 1 is excerpted therefrom. The U.S. Standard Atmosphere of Table 1 is a revised U.S. Standard Atmosphere, adapted by the United States Committee on Extension to the Standard Atmosphere (COESA) in 1962.
TABLE 1 ______________________________________ U.S. Standard Atmosphere Pressure Density Speed of Altitude Temp ratio, ratio sound V.sub.n h, ft* T, .degree.F..dagger. p/p.sub.o (.rho./.rho..sub.o) (.rho..sub.o /.rho.).sup.0.5 ft/s .dagger-dbl. ______________________________________ 0 59.00 1.0000 1.0000 1.000 1,116 5,000 41.17 0.8320 0.8617 1.077 1,097 10,000 23.34 0.6877 0.7385 1.164 1,077 15,000 5.51 0.5643 0.6292 1.261 1,057 20,000 -12.62 0.4595 0.5328 1.370 1,036 25,000 -30.15 0.3711 0.4481 1.494 1,015 30,000 -47.99 0.2970 0.3741 1.635 995 35,000 -65.82 0.2353 0.3099 1.796 973 36,089 -69.70 0.2234 0.2971 1.835 968 40,000 -69.70 0.1851 0.2462 2.016 968 45,000 -69.70 0.1455 0.1936 2.273 968 50,000 -69.70 0.1145 0.1522 2.563 968 55,000 -69.70 0.09001 0.1197 2.890 968 60,000 -69.70 0.07078 0.09414 3.259 968 65,000 -69.70 0.05566 0.07403 3.675 968 65,800 -69.70 0.05356 0.07123 3.747 968 70.000 -67.30 0.04380 0.05789 4.156 971 75,000 -64.55 0.03452 0.04532 4.697 974 80,000 -61.81 0.02725 0.03553 5.305 977 85,000 -59.07 0.02155 0.02790 5.986 981 90,000 -56.32 0.01707 0.02195 6.970 984 95,000 -53.58 0.01354 0.01730 7.600 988 100,000 -50.84 d.01076 0.01365 8.559 991 ______________________________________ * .times. 0.3048 = meters. .dagger. .times. (.degree.F. - 32)/1.8 = .degree.C. .dagger-dbl. .times. 0.3048 = m/s.
The values given up to about 65,000 feet are designated as standard. The region from 65,000 to 105,000 feet is designated proposed standard. The assumed sea-level conditions are: pressure, P.sub.o =29.91 inches (760 mm) of Hg=2,116.22 lb/ft.sup.2 ; mass density .rho..sub.O =0.002378 slugs/ft.sup.3 (0.001225 gm/cm.sup.3); T.sub.O =59.degree. F. (15.degree. C.). These assumed sea level conditions define what is known as a "standard day."
At this point, it is useful to define the term "density altitude" (H.sub.d). Density altitude is a convention by which air density is expressed in terms of the U.S. Standard Atmosphere altitude at which the corresponding air density is encountered. Referring to Table 1, when the actual air density outside the helicopter is 0.8617.times.sea level conditions, the density altitude, H.sub.d, equals 5,000 feet, and when the actual air density is 0.7385.times.sea level density, the density altitude, H.sub.a, equals 10,000 feet. It is important to note that as density altitude H.sub.d increases, air density is decreasing and when density altitude H.sub.d decreases, air density is increasing. Using relations developed from the Standard Atmosphere, air density expressed in terms of density altitude may be calculated using a term representing the pressure altitude H.sub.p, i.e., the altitude based on pressure values such as obtained from an altimeter, and a term using outside air temperature (OAT). In general, density altitude H.sub.d corresponds to the pressure altitude H.sub.p on a standard day, i.e., a day on which actual conditions match those of the standard atmosphere. Subsequent references to "altitude" in this description should be interpreted to mean "density altitude" unless the specific context indicates otherwise.
When operating helicopters at high altitudes, the lower air density reduces the tail rotor authority available to control the helicopter. For this reason, many helicopters have a lower cross wind rating at high altitude than they do at low altitude. By increasing the maximum pitch of the tail rotor blades, the full rated authority of the tail rotor could be restored at higher altitudes. However, the drive assembly and helicopter airframe must be designed to take the stress produced by the maximum possible tail rotor authority. Thus, a design which would allow a tail rotor pitch at high altitudes that produced full tail rotor authority would have to be strong enough to withstand the stresses imposed when the same pitch was used at a lower altitude in more dense air. If this higher blade angle was available at all altitudes, this would require, at a minimum, extensive testing and requalification of the helicopter airframe and transmission system, and possibly the use of a much heavier design if the tail rotor forces are found to be excessive for the existing structure, the overstressing the tail rotor gear box, tail boom, or overtorquing this tail rotor drive train.
Restated, the need for this invention can be summarized as follows. When a helicopter is designed for low altitude use, its subsequent use at higher altitudes will result in less control because less tail rotor authority is available in the less dense air. Because of the reduced control at higher altitudes, restrictions may be imposed on the use of the helicopter, such as load limits or cross wind limits. Alternately, where a helicopter is designed to have adequate control at higher altitudes, it will be over-designed for use at low altitudes. This over-design may result in higher costs or in increased weight, which in turn would reduce the helicopter's useful load carrying capacity. In addition, a helicopter going from low altitude service to high altitude service cannot be simply converted because the addition of a conventional tail rotor producing greater thrust will require the requalification of the helicopter by the appropriate governmental regulatory authorities, typically an expensive and time-consuming task. The current invention addresses these problems in the following ways: It allows the helicopter to maintain maximum tail rotor force at all altitudes; there is little or no weight penalty associated with the use of the current invention; it preserves the flight characteristics, i.e., the control "feel" and control pedal travel distances; it eliminates the need to requalify the airframe or tail rotor transmission when a more powerful tail rotor is provided; and it provides the pilot with useful information regarding ambient air conditions.
The present invention allows a helicopter to take advantage of increased thrust produced by larger or more powerful tail rotor blades by varying the maximum tail rotor pitch angle in relation to density altitude. The maximum tail rotor pitch will be controlled by the invention to limit the maximum tail rotor thrust produced by the new blades to be equivalent of the current production tail rotor blade's operation at sea level. This will limit the thrust generated to the previously qualified values and will therefore generate no excessive tail boom load or tail rotor drive shaft torque values. By making the maximum tail rotor thrust variable, the current invention allows the tail rotor to be compensated for loss of effectiveness due to increasing density altitude.
Since it is common for helicopters to be employed in different activities during their useful lives, it is not uncommon to find a helicopter originally acquired for low altitude work which is subsequently pressed into high altitude work. A need exists for an apparatus which controls the maximum pitch of the tail rotor blades in relation to density altitude and which can be added or retrofitted to existing helicopters as well as being incorporated into new helicopter designs.
Since the piloting of a helicopter requires great skill and experience, a helicopter pilot typically becomes accustomed to the control "feel", i.e., the control motions and forces needed to pilot the helicopter in a given situation. Many pilots would find modifications to the helicopter unacceptable which significantly changed the control "feel." Thus, a need exists for an apparatus allowing variable maximum pitch of the tail rotor blades in relation to density altitude but which maintains the same control pedal travel.
Since the torque produced by the turning of the helicopter's main rotor typically acts in the same direction, the needs for tail rotor authority are not symmetrical. For example, on a helicopter whose main rotor blades turn counterclockwise when viewed from above, the air frame of such helicopter will always be subject to clockwise torque. This clockwise torque will assist yaw motions to the right while opposing yaw motions to the left. It is frequently desirable for such a helicopter to have considerably more left tail rotor authority than right tail rotor authority. In the example of the helicopter whose main rotor blades rotate counterclockwise, increased left thrust is desired at higher altitudes, but increased right thrust is not needed. A need exists for an apparatus allowing variable maximum pitch of the tail rotor blades in one direction in relation to density altitude, while the maximum pitch of the blades in the other direction remains essentially unchanged.
The advantages of the current invention are as follows: it allows the helicopter to maintain maximum tail rotor force at all altitudes; there is little or no weight penalty associated with the use of the current invention; it preserves the flight characteristics, i.e., the control "feel" and control pedal travel distances; it eliminates the need to requalify the airframe or tail rotor transmission when a more powerful tail rotor is installed; and it provides the pilot with useful information regarding ambient air conditions.